Gas turbine engine

This application is a continuation-in-part application of U.S. application Ser. No. 18/824,100, filed Sep. 4, 2024, which is a continuation of U.S. application Ser. No. 17/972,720, filed Oct. 25, 2022, which issued as U.S. Pat. No. 12,104,539 on Oct. 1, 2024, all of which are hereby incorporated by reference in their entirety.

The present disclosure relates to a gas turbine engine, such as an aeronautical gas turbine engine.

A gas turbine engine generally includes a turbomachine. The turbomachine includes several engine accessories such as controllers, pumps, heat exchangers and the like that are necessary for operation. These engine accessories and engine systems may be mounted to the turbomachine.

Reference will now be made in detail to present embodiments of the disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the disclosure.

The word “exemplary” is used herein to mean “serving as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other implementations. Additionally, unless specifically identified otherwise, all embodiments described herein should be considered exemplary.

As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gas turbine engine or aircraft and refer to the normal operational attitude of the gas turbine engine or aircraft. For example, with regards to a gas turbine engine, forward refers to a position closer to an engine inlet and aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to both direct coupling, fixing, or attaching, as well as indirect coupling, fixing, or attaching through one or more intermediate components or features, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.

Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

The term “cowl” includes a housing, casing, or other structure that at least partially encases or surrounds a portion of a turbomachine or gas turbine engine.

The phrases “from X to Y” and “between X and Y” each refers to a range of values inclusive of the endpoints (i.e., refers to a range of values that includes both X and Y).

The term “propulsive efficiency” refers to an efficiency with which the energy contained in an engine's fuel is converted into kinetic energy for the vehicle incorporating the engine, to accelerate it, or to replace losses due to aerodynamic drag or gravity.

As used herein, the term “rated speed” with reference to a gas turbine engine refers to a maximum rotational speed that the gas turbine engine may achieve while operating properly. For example, the gas turbine engine may be operating at the rated speed during maximum load operations, such as during takeoff operations.

The term “standard day operating condition” refers to ambient conditions of sea level altitude, 59 degrees Fahrenheit, and 60 percent relative humidity.

A “third stream” as used herein means a non-primary air stream capable of increasing fluid energy to produce a minority of total propulsion system thrust. A pressure ratio of the third stream may be higher than that of the primary propulsion stream (e.g., a bypass or propeller driven propulsion stream). The thrust may be produced through a dedicated nozzle or through mixing of an airflow through the third stream with a primary propulsion stream or a core air stream, e.g., into a common nozzle.

In certain exemplary embodiments an operating temperature of the airflow through the third stream may be less than a maximum compressor discharge temperature for the engine, and more specifically may be less than 350 degrees Fahrenheit (such as less than 300 degrees Fahrenheit, such as less than 250 degrees Fahrenheit, such as less than 200 degrees Fahrenheit, and at least as great as an ambient temperature). In certain exemplary embodiments these operating temperatures may facilitate heat transfer to or from the airflow through the third stream and a separate fluid stream. Further, in certain exemplary embodiments, the airflow through the third stream may contribute less than 50% of the total engine thrust (and at least, e.g., 2% of the total engine thrust) at a takeoff condition, or more particularly while operating at a rated takeoff power at sea level, at a static flight speed, and/or at 86 degree Fahrenheit ambient temperature operating conditions.

Furthermore in certain exemplary embodiments, aspects of the airflow through the third stream (e.g., airstream, mixing, or exhaust properties), and thereby the aforementioned exemplary percent contribution to total thrust, may passively adjust during engine operation or be modified purposefully through use of engine control features (such as fuel flow, electric machine power, variable stators, variable inlet guide vanes, valves, variable exhaust geometry, or fluidic features) to adjust or optimize overall system performance across a broad range of potential operating conditions.

Conventional turbofan engine design practice has been to provide an outer nacelle surrounding the fan to provide relatively efficient thrust for the turbofan engine at high fan speeds (compared with an unducted fan). Such a configuration may generally limit a permissible size of the fan (i.e., a diameter of the fan). Generally, a turbofan engine includes a fan to provide a desired amount of thrust without overloading the fan blades (i.e., without increasing a disk loading of the fan blades of the fan beyond a certain threshold), and therefore to maintain a desired overall propulsive efficiency for the turbofan engine. The inventors of the present disclosure seek to drive the fan diameter higher, thereby to reduce fan pressure ratio while maintaining the same level of thrust to improve fuel efficiency. By increasing the fan diameter, however, an installation of the turbofan engine becomes more difficult. In addition, if an outer nacelle is maintained, the outer nacelle may become weight prohibitive with some larger diameter fans.

The inventors of the present disclosure found that for a three-stream gas turbine engine having an unducted primary fan (the outer nacelle removed) and a ducted secondary fan, with the secondary fan providing an airflow to a third stream of the gas turbine engine, an overall propulsive efficiency of the gas turbine engine that results from providing a high diameter fan may be maintained at a high level, while reducing the size of the primary fan. Such a configuration may maintain a desired overall propulsive efficiently for the gas turbine engine, or unexpectedly, may in fact increase the overall propulsive efficiency of the gas turbine engine. Further, by including a third stream, an axial length of the core engine may be reduced relative to the overall engine axial length by allowing for a portion of the airflow through the engine to flow through the third stream. This reduces an overall weight of the engine. However, the core engine must maintain a sufficient size to produce enough power to drive the primary fan and the ducted secondary fan.

Further, removing the outer nacelle and reducing the overall axial length of the core engine significantly reduces engine accessory storage space. A diameter of a core cowl may be increased to make room for the accessories between an engine casing and an inner surface of the core cowl, however, the core cowl diameter cannot be too large due to potential performance penalties such as excessive drag and installation difficulties.

The inventors proceeded in the manner of designing a gas turbine engine with a given core cowl diameter, core diameter, core axial length, and overall engine axial length; checking the propulsive efficiency of the designed gas turbine engine; redesigning the gas turbine engine with varying core cowl diameters, core diameters, core axial lengths, and overall engine axial lengths; rechecking the propulsive efficiency of the redesigned gas turbine engine; and then making accommodations when, for example, it was found that subsystem sizes increased due to certification requirements and/or power requirements, or servicing needs impacted where to locate things during the design of several different types of gas turbine engines, including the gas turbine engine described below with reference to, e.g.,.

During the course of this practice of studying and evaluating various cowl diameters, core diameters, core length, and engine length considered feasible for best satisfying mission requirements, it was discovered that certain relationships exist between a core cowl diameter ratio (which is equal to a peak cowl diameter divided by a maximum combustor casing diameter) and a core cowl length ratio (which is equal to an under-core cowl axial length divided by an overall core axial length). In particular, the inventors of the present disclosure have found that these ratios can be thought of as an indicator of the ability of a gas turbine engine to provide sufficient packaging space between the core engine combustor casing and the core cowl for packaging/mounting various accessories and/or engine systems, while also having a core engine capable of producing sufficient power to drive primary and secondary fans, particularly in more complex engine designs. In some embodiments, the inventors found that selectively coupling one or more engine components such as an engine accessory or system component to one of the core cowl or to the engine improves accessibility for inspection, repair, and maintenance and improves weigh loads on the core engine.

Referring now to the drawings,is a perspective view of an exemplary aircraftthat may incorporate at least one exemplary embodiment of the present disclosure. As shown in, the aircrafthas a fuselage, wingsattached to the fuselage, and an empennage. The aircraftfurther includes a propulsion systemthat produces a propulsive thrust to propel the aircraftin flight, during taxiing operations, etc. Although the propulsion systemis shown attached to the wing(s), in other embodiments it may additionally or alternatively include one or more aspects coupled to other parts of the aircraft, such as, for example, the empennage, the fuselage, or both. The propulsion systemincludes at least one engine. In the exemplary embodiment shown, the aircraftincludes a pair of gas turbine engines. Each gas turbine engineis mounted to the aircraftin an under-wing configuration. Each gas turbine engineis capable of selectively generating a propulsive thrust for the aircraft. The gas turbine enginesmay be configured to burn various forms of fuel including, but not limited to unless otherwise provided, jet fuel/aviation turbine fuel, and hydrogen fuel.

is a cross-sectional side view of a gas turbine enginein accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of, the gas turbine engineis a multi-spool, high-bypass turbofan jet engine, sometimes also referred to as a “turbofan engine.” As shown in, the gas turbine enginedefines an axial direction A (extending parallel to a longitudinal centerlineprovided for reference), a radial direction R, and a circumferential direction C extending about the longitudinal centerline. In general, the gas turbine engineincludes a fan sectionand a turbomachinedisposed downstream from the fan section.

The exemplary turbomachinedepicted generally includes an engine housing, casing, or core cowlthat defines an annular core inlet. The core cowlat least partially encases, in serial flow relationship, a compressor section including a booster or low-pressure compressorand a high-pressure compressor, a combustion section, a turbine section including a high-pressure turbineand a low-pressure turbine, and at least a portion of a jet exhaust nozzle. Together, these components or sections make up a core engineof the turbomachine.

A high-pressure shaftdrivingly connects the high-pressure turbineto the high-pressure compressor. A low-pressure shaftdrivingly connects the low-pressure turbineto the low-pressure compressor. The compressor section, combustion section, turbine section, and jet exhaust nozzletogether define a working gas flow paththrough the gas turbine engine.

For the embodiment depicted, the fan sectionincludes a fanhaving a plurality of fan bladescoupled to a diskin a spaced apart manner. As depicted, the fan bladesextend outwardly from diskgenerally along the radial direction R. Each fan bladeis rotatable with the diskabout a pitch axis P by virtue of the fan bladesbeing operatively coupled to a suitable pitch change mechanismconfigured to collectively vary the pitch of the fan blades, e.g., in unison. The fan blades, disk, and pitch change mechanismare together rotatable about the longitudinal centerlineby the low-pressure shaft.

In an exemplary embodiment, as shown in, the gas turbine enginefurther includes a power gearbox or gearbox. The gearboxincludes a plurality of gears for adjusting a rotational speed of the fanrelative to a rotational speed of the low-pressure shaft, such that the fanand the low-pressure shaftmay rotate at more efficient relative speeds. The gearboxmay be any type of gearbox suitable to facilitate coupling the low-pressure shaftto the fanwhile allowing each of the low-pressure turbineand the fanto operate at a desired speed. For example, in some embodiments, the gearboxmay be a reduction gearbox. Utilizing a reduction gearbox may enable the comparatively higher speed operation of the low-pressure turbinewhile maintaining fan speeds sufficient to provide for increased air bypass ratios, thereby allowing for efficient operation of the gas turbine engine. Moreover, utilizing a reduction gearbox may allow for a reduction in turbine stages that would otherwise be present (e.g., in direct drive engine configurations), thereby providing a reduction in weight and complexity of the engine.

Referring still to the exemplary embodiment of, the diskis connected to the gearboxvia a fan shaft. The diskis covered by a rotatable front hubof the fan section(sometimes also referred to as a “spinner”). The front hubis aerodynamically contoured to promote an airflow through the plurality of fan blades. Additionally, the exemplary fan sectionincludes an annular fan casing or outer nacellethat circumferentially surrounds the fanand/or at least a portion of the turbomachine. The nacelleis supported relative to the turbomachineby a plurality of circumferentially spaced struts or outlet guide vanesin the embodiment depicted. Moreover, a downstream sectionof the nacelleextends over an outer portion of the turbomachineto define a bypass airflow passagetherebetween.

is a schematic cross-sectional view of a portion of the core engineof the gas turbine engineas shown in, according to an exemplary embodiment of the present disclosure. As shown in, the high-pressure compressoris encased within a compressor casing. The combustion sectionis encased within a combustor casing. The high-pressure turbineand the low-pressure turbineare encased within one or more turbine casing(s). The combustor casingdefines an outer surface. A void or spaceis defined between an inner surfaceof the core cowland the outer surfaceof the combustor casing. The core cowlfurther includes an outer surfaceradially spaced from the inner surfacewith respect to radial direction R. In exemplary embodiments, at least one engine componentis coupled to the core cowlinner surface. The at least one engine componentmay include but is not limited to valves, electronics including engine and system controllers, fire and overheat detection system components, fire extinguisher components, heat exchangers, pumps, generator, etc.

In exemplary embodiments, engine componentis selectively coupled to the core engineor the core cowl. When the engine componentis coupled to the core cowl, the engine componenttravels with the core cowlwhen pivoted away from the core engine. When the engine componentis coupled to the core engine, the engine componentstays coupled to the core enginewhen the core cowlis pivoted away from the core engine. In exemplary embodiments and as previously presented, the engine componentis one of a heat exchanger, a sensor, a controller, a pump, a duct, a valve, fire and overheat detection system components, fire extinguisher components, or a generator. It should be appreciated that this list is not all inclusive of possible engine components that may be selectively coupled to the core cowlor the core engine.

In exemplary embodiments, the engine componentis selectively coupled to the core engineor the core cowlvia a fastener. As shown in, the fastenermay be disposed between a core cowl structuresuch as a strut or bracket, and a core engine structuresuch as a strut, a casing or bracket. The core cowl structuremay be fixedly coupled to the core cowl, such that the core cowl structuremoves with the core cowl, as described below. By contrast, the core engine structureis not moveable with the core cowland instead may be fixedly coupled to a stationary and structural component of the core engine, such as the compressor casing(as in the embodiment depicted), or one or more of the combustor casing, turbine casing, or a support frame such as a compressor frame, a mid-frame, or a rear support frame or turbine frame, etc.

The fastenermay be fixedly connected to the engine component. The fastenermay comprise a cam lock type fitting, bayonet fitting, quarter-turn fastener or other mechanical or electromechanical fastener or device that allows selectively coupling the engine componentto the core cowlor the core engine. In particular embodiments, the core cowldefines or includes an access opening or hatchwherein the fasteneris accessible from the access opening.

It should be appreciated, however, that the exemplary gas turbine enginedepicted inis provided by way of example only, and that in other exemplary embodiments, the gas turbine enginemay have other configurations. For example,is a schematic cross-sectional view of a gas turbine engineaccording to another example embodiment of the present disclosure. Particularly,provides a turbofan engine having a rotor assembly with a single stage of unducted rotor blades. In such a manner, the rotor assembly may be referred to herein as an “unducted fan,” or the entire enginemay be referred to as an “unducted turbofan engine.” In addition, the engineofincludes a third stream extending from the compressor section to a rotor assembly flowpath over the turbomachine, as will be explained in more detail below.

For reference, the enginedefines an axial direction A, a radial direction R, and a circumferential direction C. Moreover, the enginedefines an axial centerline or longitudinal axisthat extends along the axial direction A. In general, the axial direction A extends parallel to the longitudinal axis, the radial direction R extends outward from and inward to the longitudinal axisin a direction orthogonal to the axial direction A, and the circumferential direction extends three hundred sixty degrees (360°) around the longitudinal axis. The engineextends between a forward endand an aft end, e.g., along the axial direction A.

As shown inthe engineincludes a turbomachinehaving a fan sectionthat is positioned upstream thereof. Generally, the turbomachineincludes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. Particularly, as shown in, the turbomachineincludes a housing or core cowlthat defines an annular core inlet. The core cowlfurther encloses at least in part a low-pressure system and a high-pressure system. For example, the core cowldepicted encloses and supports at least in part a booster or low-pressure (“LP”) compressorfor pressurizing the air that enters the turbomachinethrough core inlet. A high-pressure (“HP”), multi-stage, axial-flow compressorreceives pressurized air from the LP compressorand further increases the pressure of the air. The pressurized air stream flows downstream to a combustorof the combustion section where fuel is injected into the pressurized air stream and ignited to raise the temperature and energy level of the pressurized air.

It will be appreciated that as used herein, the terms “high/low speed” and “high/low-pressure” are used with respect to the high-pressure/high speed system and low-pressure/low speed system interchangeably. Further, it will be appreciated that the terms “high” and “low” are used in this same context to distinguish the two systems and are not meant to imply any absolute speed and/or pressure values.

The high energy combustion products flow from the combustordownstream to a high-pressure turbine. The high-pressure turbinedrives the high-pressure compressorthrough a high-pressure shaft. In this regard, the high-pressure turbineis drivingly coupled with the high-pressure compressor. The high energy combustion products then flow to a low-pressure turbine. The low-pressure turbinedrives the low-pressure compressorand components of the fan sectionthrough a low-pressure shaft. In this regard, the low-pressure turbineis drivingly coupled with the low-pressure compressorand components of the fan section. The low-pressure shaftis coaxial with the high-pressure shaftin this example embodiment. After driving each of the high-pressure turbineand the low-pressure turbine, the combustion products exit the turbomachinethrough a rear support frame or turbomachine exhaust nozzle. A core engineof the gas turbine engineis defined as the part of the gas turbine enginethat extends from the fan sectionto the rear support frame or turbomachine exhaust nozzle.

Accordingly, the turbomachinedefines a working gas flowpath or core ductthat extends between the core inletand the rear support frame or turbomachine exhaust nozzle. The core ductis an annular duct positioned generally inward of the core cowlalong the radial direction R. The core duct(e.g., the working gas flowpath through the turbomachine) may be referred to as a second stream. The fan sectionincludes a fan, which is the primary fan in this example embodiment. For the depicted embodiment of, the fanis an open rotor or unducted fan. In such a manner, the enginemay be referred to as an open rotor engine. Moreover, it will be appreciated that the fan sectionincludes a single fan, and the fanis the only unducted fan of the gas turbine enginedepicted.

As depicted, the fanincludes a plurality or an array of fan blades(only one shown in). The fan bladesare rotatable, e.g., about the longitudinal axis. As noted above, the fanis drivingly coupled with the low-pressure turbinevia the low-pressure shaft. For the embodiments shown in, the fanis coupled with the low-pressure shaftvia a speed reduction gearbox, e.g., in an indirect-drive or geared-drive configuration.

Moreover, the array of fan bladescan be arranged in equal spacing around the longitudinal axis. Each fan bladehas a root and a tip and a span defined therebetween. Each fan bladedefines a central blade axis. For this embodiment, each fan bladeof the fanis rotatable about its central blade axis, e.g., in unison with one another. One or more actuatorsare provided to facilitate such rotation and therefore may be used to change a pitch of the fan bladesabout their respective central blades' axes.

The fan sectionfurther includes a fan guide vane arraythat includes fan guide vanes(only one shown in) disposed around the longitudinal axis. For this embodiment, the fan guide vanesare not rotatable about the longitudinal axis. Each fan guide vanehas a root and a tip and a span defined therebetween. The fan guide vanesmay be unshrouded as shown inor, alternatively, may be shrouded, e.g., by an annular shroud spaced outward from the tips of the fan guide vanesalong the radial direction R or attached to the fan guide vanes.

Each fan guide vanedefines a central blade axis. For this embodiment, each fan guide vaneof the fan guide vane arrayis rotatable about its respective central blade axis, e.g., in unison with one another. One or more actuatorsare provided to facilitate such rotation and therefore may be used to change a pitch of the fan guide vaneabout its respective central blade axis. However, in other embodiments, each fan guide vanemay be fixed or unable to be pitched about its central blade axis. The fan guide vanesare mounted to a fan housing or fan cowl.

As shown in, in addition to the fan, which is unducted, a ducted fanis included aft of the fan, such that the engineincludes both a ducted and an unducted fan which both serve to generate thrust through the movement of air without passage through at least a portion of the turbomachine(e.g., without passage through the HP compressorand combustion section for the embodiment depicted). The ducted fanis rotatable about the same axis (e.g., the longitudinal axis) as the fan blade. The ducted fanis, for the embodiment depicted, driven by the low-pressure turbine(e.g., coupled to the low-pressure shaft). In the embodiment depicted, as noted above, the fanmay be referred to as the primary fan, and the ducted fanmay be referred to as a secondary fan. It will be appreciated that these terms “primary” and “secondary” are terms of convenience, and do not imply any particular importance, power, or the like.

The ducted fanincludes a plurality of fan blades (not separately labeled in) arranged in a single stage, such that the ducted fanmay be referred to as a single stage fan. The fan blades of the ducted fancan be arranged in equal circumferential spacing around the longitudinal axis. Each blade of the ducted fanhas a root and a tip and a span defined therebetween.

The fan cowlannularly encases at least a portion of the core cowland is generally positioned outward of at least a portion of the core cowlalong the radial direction R. Particularly, a downstream section of the fan cowlextends over a forward portion of the core cowlto define a fan duct flowpath, or simply a fan duct. According to this embodiment, the fan flowpath or fan ductmay be understood as forming at least a portion of the third stream of the engine.

Incoming air may enter through the fan ductthrough a fan duct inletand may exit through a fan exhaust nozzleto produce propulsive thrust. The fan ductis an annular duct positioned generally outward of the core ductalong the radial direction R. The fan cowland the core cowlare connected together and supported by a plurality of substantially radially extending and circumferentially spaced stationary struts(only one shown in).

The stationary strutsmay each be aerodynamically contoured to direct air flowing thereby. Other struts in addition to the stationary strutsmay be used to connect and support the fan cowland/or core cowl. In many embodiments, the fan ductand the core ductmay at least partially co-extend (generally axially) on opposite sides (e.g., opposite radial sides) of the core cowl. For example, the fan ductand the core ductmay each extend directly from a leading edgeof the core cowland may partially co-extend generally axially on opposite radial sides of the core cowl.